Rocket means for driving a free punch

ABSTRACT

ROCKET LAUNCHING AND DRIVE MEANS FOR DRIVING A FREE PUNCH BY A ROCKET THAT IS MOUNTED INSIDE A CYLINDRICAL LAUNCHING TUBE BY SEGMENTED SABOT MEANS AND HAS A SPIN TURBINE SECURED TO THE ROCKET TO ROTATE THE ROCKET AS IT IS BEING LAUNCHED. THE SEGMENTED SABOT MEANS AND SPIN TURBING ARE SECURED TO THE ROCKET SO AS TO BE SEPARATED FROM THE ROCKET BY CENTRIFUGAL FORCE AS THE ROCKET LEAVES THE LAUNCHING TUBE.

Oct. 5, 1971 v. BLACK ETAL 3,610,095

nocxmw MEANS FOR DRIVING FREE PUNCH Original Filed May 18, 1962 3 Sheets-Sheet l V. Black INVENTORS.

Donald Albert D. Jomtuos Ross T; Rodey John K. Wall,

o. v. BLACK ETA!- 3,610,095

ROCKET MEANS FOR DRIVING A FREE PUNCH 3 Sheets-Sheet 8 Original Filed May 18, 1962 BUFINOUT RANGE(FEET) FIG.6

Donald V. Block Albert D. Jomtoos Ross T. Rodey John KWoH,

INV NTORS.

Oct. 5, 1971 v, CK ET AL ROCKET MEANS FOR DRIVING A FREE PUNCH 3 Sheets-Sheet 5 Original Filed May 18. 1962 V=7000FPS FOR 7687? PENETRATION I2" LEAD ARBAL|ST% TUNGSTEN URANIUM 238 ESTIMATED GROSS WTOF ROCKET,

1o SLUG DENSlTY(LB/|N |o VELOClTYHOOOFT/SEC) FIG. 7

n 0 2 :5 -m I 9 H 8 Wm -7 E\A D 6 1 5 /fr 4 M X c w i. W w m A l- 2 w R w m n I I l h I I i m9a7 5 4 3 2 ARMOR THICKNESS l CORE DIAMETER d mv NTORS.

FIG. 8

United States Patent Patented Oct. 5., 1971 ROCKET MEANS FOR DRIVING A FREE PUNCH Donald V. Black, Santa Monica, Albert D. Jamtaas, Los Angeles, Ross T. Radey, Palos Verdes, and John K. Wall, Los Angeles, Calif., assignors to the United States of America as represented by the Secretary of the Army Application Aug. 10, 1965, Ser. No. 480,248, which is a division of application Ser. No. 196,854, May 18, 1962. Divided and this application June 29, 1967, Ser. No.

Int. Cl. F41f 3/04 US. Cl. 891.8 Claims ABSTRACT OF THE DISCLOSURE Rocket launching and drive means for driving a free punch by a rocket that is mounted inside a cylindrical launching tube by segmented sabot means and has a spin turbine secured to the rocket to rotate the rocket as it is being launched. The segmented sabot means and spin turbine are secured to the rocket so as to be separated from the rocket by centrifugal force as the rocket leaves the launching tube.

This application is a division of application Ser. No. 480,248, filed Aug. 10, 1965, now US. Pat. -No. 3,547,031, which is a division of application Ser. No. 196,854, filed May 18, 1962.

BACKGROUND OF THE INVENTION The present invention relates to a device for punching holes in massive structures such as heavy armor, spaced or continuous, having a total thickness of six to twelve inches, for example, and relate more particularly to a missile drive system and the components thereof. Such armor is found on naval craft and on land vehicles such as tanks. Many devices have been tried for disabling tanks, such as land mines, artillery and bazookas, for example, but such devices have not been fully successful.

There is need for a device which will punch a hole in any armor which a tank or ship could conceivably carry and at the same time not be foiled as by false coverings which either deflect or trigger premature action, or foiled by layers of spaced armor plate. And there is need for a punch which can be launched from and be effective at distances of a quarter to about two miles from the target. Further, the launching of the punch must be achieved by, preferably, one man aiming and firing equipment which can be easily carried by him. Visual sighting of the launcher must be simple and quick. The weight of a complete launcher, power plant, and punch should not exceed twenty pounds if the device is to fulfill such requirements. This performance ability with weight limitations is not found in the prior art. Also, there is need for a punch for use in attacking concrete and rock. The concrete may be either reinforced with steel rods or steel aggregate, or nnreinforced with such materials. In fact, there is need for a punch which will effect penetration of a wide variety of materials for a wide variety of purposes. These purposes occur both under military and under civilian conditions.

It is imperative to the utility of a missile that it has a very high probability of a first-round hit. Until very recently, this requirement could not be met by highvelocity guns, except, of course, at extremely short ranges, point-blank, a condition where unguided low velocity rockets carrying shaped charges are fairly adequate. However, guns large enough to be lethal to a heavy tank are themselves very heavy and expensive. As a result, adequate antitank gun support is rarely available to troops under tank attack. Recently, small guided missiles of low velocity and using shaped charges have become available. These missiles have the advantage over guns of only requiring light and portable launchers, but, primarily, due to their guidance means, they are relatively unreliable and expensive.

The present invention relates to a rocket and punch which fills this need for an antitank weapon, a weapon that can be handled and fired by a single man, that has an eifective range of a quarter to two miles, and that has a high probability of first-round hit and kill. The present invention, also, relates to a free punch which can be launched remotely from a target, will have a high degree of accuracy as to hitting the target just where desired, and which will penetrate through or deeply into any material.

Thus, it will be seen that it is an object of the present invention to devise a system for punching holes in massive objects such as thick armor plate of means of a punch which is launched remotely from the target and which systems components are very light in weight, and which punch can be put on target by a simple and direct sight because of a very short flight time.

Further objects of the present invention result from fundamental aspects of the present device. If the weight of the device is to be within the given limit, the punch weight must be close to or under one pound, the power plant must be under twelve pounds, and the cartridgelauncher must not exceed ten pounds. The only manner in which a free one-pound punch can penetrate twelve inches of armor is for the punch to have a density, a slenderness, and a velocity that will give a high enough impact momentum per unit area of punch cross section or unit area of punch and target contact to be effective. This means a critical proportioning of weight, power, and dimensions. The weight and dimensions of a launcher and a rocket motor to drive a punch are critical if mobility is to be achieved with needed punch weight and velocity. Thus, the power available is critical. The power available dictates the weight of the punch, and this dictates the density and slenderness of the punch necessary to achieve the unit impact pressures required to effect penetration. The prior art does not teach such proportioning or how to achieve such unit pressures. Further, the art does not teach how to deliver such a free punch in proper aspect, so that its longitudinal axis is tangent to its flight path.

Thus, it becomes a further object of the invention to provide a punch and rocket drive therefor which may be transported and launched by a single person such as a foot soldier, and which punch and rocket, and punch, rocket, and launcher are so proportioned to each other and to their components that such proportioning is critical and critical to the resulting punch function.

Further objects relating to the carrying out of the primary objects are the design of a rocket case which is light enough and strong enough so that the primary objects may be achieved, and which case is strong enough so that the ratio of easing weight to grain Weight is smaller than formerly thought possible; and the proportion of weight to punch velocity is more favorable than formerly thought possible; the design of a turbine which will spin the rocket and punch sufliciently to insure a highly accurate flight, and which will drop off at the end of the launching; the design of a connector between rocket casing and punch which will hold them together in flight, which allows easy punch exchangeability, and which permits and induces desirable energy exchange between rocket and punch; the design of sabots which will drop off at launch and which will minimize tip-01f; the design of a rocket and launch system so that for a given range there will be self-compensation in the system for crosswinds; and the design of a rocket so that aiming 3 errors will be practically eliminated due to the velocity of its flight, its time of flight being of the order of one second.

SUMMARY OF THE INVENTION The aforementioned objects and others herein apparent are achieved by a rocket having in its nose cone a cylindrical punch, or slug, with its axis coaxial of the cone and the body of the rocket; which slug is slender and weighs about a pound, and is made of a dense material such as tungsten or an alloy thereof, or depleted uranium-238, or alloy thereof, and which rocket has a gross weight of about ten pounds, ready to launch but exclusive of its launcher, with a capability of achieving a speed of 8,000 feet per second in 0.6 second; and with a spin turbine, which drops off at launch, to give the rocket accuracy in flight. Such a rocket may be launched in a manner similar to a bazooka and with similar equipment but will have an effective range of about 1,200 to 10,000 feet for heavy armor penetration. Such a rocket will have a very flat trajectory and may be sighted with a conventional, optical system with little lead and elevation.

BRIEF DESCRIPTION OF THE DRAWING Hereinafter, there is described in detail a rocket device conforming to the above outline and capabilities, which achieves and will achieve the aforementioned objects, and which is illustrated in the drawing herewith in which:

FIG. 1 is an elevational view of a rocket embodying the present invention, which rocket is shown in its launching tube shown in longitudinal section, and which rocket is shown with its firing means ready for firing.

FIG. 2 is a longitudinal sectional view of the rocket and launching tube of FIG. 1 but without the firing mechanism, and With parts of the rocket casing in full view.

FIG. 3 is a sectional view of the line 33 of FIG. 1.

FIG. 4 is a detailed isometric view of a portion of the firing means and of a portion of the rocket case to which such means are to be attached.

FIG. 5 is a detailed isometric exploded view of a spin turbine and the locking sections which aid in securing the turbine to the back end of the rocket, shown fragmentally.

FIG. 6 is a graph of various operating characteristics of the rocket against rocket travel distance.

FIG. 7 is a graph of estimated gross weights of rockets against rocket velocities, and punch densities and slenderness.

FIG. 8 is a graph of ballistic limits against the ratios of penetrations to punch diameters.

DESCRIPTION OF THE PREFERRED EMBODIMENT A rocket assembly or vehicle 11 embodying the present invention is shown in FIG. 2 with parts thereof cut away to illustrate the internal construction thereof. There are three main sections of the rocket: a flared nozzle section 12, a fuel containing body section 13, and a nose section 14 which contains a punch or punch member 16. The body or fuel case 13 must be designed to withstand pressures of 3,500 to 4,000 p.s.i. during burning of the fuel contained therein. In closed end thin wall pressure vessels, the longitudinal stresses are one-half the circumferential stresses. This means that in a pressure cylinder made of homogenous material, the material is understressed longitudinally. The present construction materially reduces the weight of the rocket body by using an aluminum liner 17 which is only thick enough to carry the longitudinal stresses and about half of the circumferential stresses. The other half of the circumferential stresses are carried by a circumferential wrapping 18, or winding, of high strength fiberglass cords which are stabilized, held in place and protected against abrasion, by impregnation of the Wrap with an epoxy resin and curing agent. The tension of the wrap is such that the aluminum liner, under pressure, is allowed to develop its full tensile strength circumferentially. Another function of the aluminum liner is that of sealing the Wrap formed by the glass cords and resin against radial gas leakage under pressure. A still further aspect of the use of an aluminum liner is that it may be anodized so as to surface it with an aluminum oxide coating which is good refractory and is resistant to the temperatures and abrasion to which the inside surfaces of the liner are subjected during the burning of the contained solid fuel 19.

A rocket designed in accordance with the present disclosure has an over-all length of 41.20", a body outside diameter of 2.60, a nose cone length of 10.0 and a nozzle length of 5.70" with an outside diameter of 3.56". With these dimensions, if the liner thickness is 0.032" and the wrap thickness is 0.017 with respective densities of 0.1 and 0.07 lb./in. there will be a saving of almost a pound in weight of the body as compared with the use of aluminum only (double the thickness of the liner) in the body.

The liner material is carried aft of the body section and flared to provide a covering 21 for a ceramic nozzle 22. The outside of the nozzle sleeve is in the general shape of a truncated cone fitted and secured to the inside of the cover 21. The after edge portion of the nozzle extends to about the edge of the covering 21 which is adapted to butt a portion of a spin turbine. Adjacent this edge, the cover 21 is formed with an external peripheral groove 23 that cooperates with such turbine. The throat 24 of the nozzle is in its forward part, and the inside of the nozzle is flared outwardly, both forwardly and rearwardly from its throat. The use of ceramic protects the nozzle 22 from the high heat loads imposed by the exhaust.

The forward end, the nose end, of the body is closed by a heavy steel or aluminum nose disc 26, skirted and domed, that is convex forwardly. The after outside portion 27 of the skirt is formed parallel to and contiguous with the inside of the liner 17 at its forward end, and thereat the liner and the disc are welded together. The forward portion 28 of the skirt is threaded to receive thereover and retain the base portion of a nose cover cone 34 which provides fairing for the punch 16. The voids inside of the nose cone may be filled by a foamed-in-place plastic 35 to reinforce the cone so that it may be formed of light Weight material.

The punch is, for example, a solid cylinder 7.77" long and 0.52" thick to give a slenderness ratio of 15. It is made of tungsten or uranium or an alloy of either material. The density will be of the order of 0.53 to 0.66 lb./ in. The aft end of the punch is threaded to screw into a socket 36 located centrally of the nose disc 26 so as to hold the punch with longitudinal axis coaxial of the rocket. The threading of the punch to the disc 26 makes for easy assembly and for change of punches, depending on the type of target.

The above details describe the net, or empty, rocket. Added to this to give the gross, or full, rocket is the solid propellant fuel 19, or grain, and its igniter 38 with its fuze pigtail 39. The igniter 38 is nozzle-throat mounted for ease of installation and to serve as a weather seal at the throat 24 for the grain under ready conditions. It is full grain-length to achieve full ignition with low time delay and is center supported by consumable foam annuli 40. There may, also, be included in the gross of the rocket, a spin turbine 41 that is secured aft of the nozzle 22 but drops off after launching of the rocket. Many considerations go into the selection of a suitable fuel, and these will not be discussed in detail here as such considerations are well known to rocket designers. It is sufficient to state that the fuel will be of the solid type such as Omax S-2b which is cast in place in the rocket body. Omax S2b is a product of Olin Mathison Chemical Corporation. The grain 19 will have a cross section as shown in FIG. 3'. The grain should have a burning time of about 0.3 to 0.5 second at 120 to -40 F. and develop pressures between 3,500 and 2,000 psi. at substantially constant pressure during burning.

The rocket may be shipped in a container which, also, serves as a launching tube, and such will be here described. The previously described rocket is shown in such a tube in both FIGS. 1 and 2 which are elevational views with parts sectioned and broken away for clarity of detail. This launching tube 42, shown in longitudinal section in both views, may be constructed in the same manner as the body 13 of the rocket by the use of an anodized tubular cylinder for the liner and a stabilized wrap of fiberglass cords and cured resins, but it is here shown as a plain aluminum tube. Each of the open ends of the tube is selectively closable by means of a friction fitting closure 46 shown only in FIG. 2. The length of the tube 42 is slightly more than that of the rocket. The after edge of the launch tube, the edge adjacent the rocket nozzle, is formed with a small slot 47 by means of which there is, at the time of firing, secured inside the tube a percussion type firing cap 48 that has attached thereto one end of the detonator pigtail 39. When the after end opening is closed by one of the closures 46, its skirt covers the slot 47 against the entrance of moisture and dirt. This tube may serve as a cartridge case or as a launching tube when it is auxiliary equipped with trigger mechanism, sight, and holding or support means.

The illustrated auxiliary launching equipment of FIG. 1 is only that necessary for a shoulder supported launch. This equipment may be secured to the tube directly, or to spaced apart straps 49', each circumferential of the tube 42, as illustrated. A shoulder rest 5 is secured to the straps 49 just forward of the center of gravity of the assembly, and forward of the rest is a pistol grip shaped hand hold 51. The pistol grip 51 carries a trigger 54 that is connected, or extended, by covered cable 56, Bowden wire, to a double acting trigger mechanism and firing pin 57 actuated thereby and contained in a firing block 58 which is placed in opposition to the firing cap 48 so that first actuation of the trigger will cause the pin 57 to cock, and a second actuation will release the pin to impinge the cap 48 to cause its detonation and initiation of the burning of the pigtail 39 to the rocket fuel igniter 38. Firing pin actuation is an old art, and the details of the present actuation are not shown. The firing block 58 is releasably secured in a slot 59 formed in the after edge of the tube adjacent the firing cap slot 47, and the cap and firing block are so aligned when in their respective slots 47 and 59 that upon actuation of the firing pin, it will strike and detonate the cap. This arrangement permits easy attachment and disengagement of the cap and firing block to and from the launching tube when the rear closure 46 is removed.

Another piece of equipment used in launching the rocket is a sabot 61. The one herein illustrated and described in full length is made in three sections, 62, 63, and 64 or petals which extend longitudinally of the rocket and circumjacent the body thereof between body and launcher tube 42 and a portion of the nose cone 34. The rocket rides on and is guided by the sabot as it travels through the tube in its launch. That is, the rocket is supported by the sabot and allowed to rotate relative to launcher tube 42 as the rocket is launched. The sabot petals separate and fall from the rocket as it, and they leave the tube. The sabot is made of balsa wood or other light material.

The last piece of equipment needed for the launching of the rocket is the spin turbine 41. This turbine comprises a one-inch long cylindrical ring 66 having the same inside diameter as the inside of the after edge of the nozzle 22, with a plurality of blades 67, about twenty, secured inside of and to the ring. Each blade is set at thirty degrees with the center line of the rocket and protrudes three-fourths inches centripetally of the ring. The forward edge of the ring is flanged outwardly and butts the rear edge of the nozzle cover 21. Also, the butting flange of the ring forms an exterior groove 68 similar to, close to, and in axial alignment with the groove 23 in the nozzle cover 21. Also, the rear edge of the nozzle is formed with three equally circumferentially spaced lugs 69 that extend forward through notches 70 in the butting edge flange of the nozzle. Three pieces of locking ring sections 71, each having a U-shaped cross section, encompass these circumferential edges, one leg of the U-shape being in one groove 23 and the other leg in the other groove 68. Each ring section is between two of the lugs 69. These ring sections 71 are retained in place by the inside of the launch tube 42 when the rocket is therein, and the rings fall away or are forced by centrifugal force from the rocket upon the rocket and turbine leaving the launch tube. The lugs 69 serve to prevent rotation of the turbine with respect to the rocket and to center the turbine coaxially thereof by the lugs contacting the outside of the nozzle. The turbine is formed of cast stainless steel. The turbine blades are not shown in FIG. 1.

The design and fueling of the rocket are such that in flight it will have a very flat trajectory whose apogee is less than four feet for target ranges of 1.25 miles. For this range, small changes in missile speed have a negligible effect on the impact point. For example, a one percent change in the burn-out speed results in only a six-inch vertical change in the impact point. If probable deviations from the nominal of drag, air density, wind velocity, impulse, and weight are all present, the probable vertical error would be less than a foot for such range. While horizontal and vertical dispersion is important for a rocket of this type, due to the rockets small size, light weight, long and slender shape, rotation, and high acceleration, it is possible to design it in such a way as to minimize the sources of such errors. Because of the rockets small size and weight, it can be launched from a light and mobile tube and aimed by direct sight. From firing, the rocket will travel the first mile in 0.95 second so that at this range, the gunner need lead a tank traveling at thirty miles per hour by only about half the length of the tank.

The rocket is launched from the tube 42 with a close fit between the rocket and tube being maintained by the use of the sabot assembly 61. The rocket is guided for about three feet of travel before the turbine 41 passes out of the tube. During this period, the rocket is accelerated to a roll rate of 800 radians per second by the reaction type spin turbine 41 attached to the rear of the nozzle 22. After leaving the launch tube, the spin turbine 41 and the sabot petals, 6-2, 63, and '64 are separated from the rocket by the action of the centrifugal force and rocket acceleration. When the turbine leaves the tube, the inner wall of the tube no longer holds the locking sections 71 in place to retain the turbine on the open end of the nozzle cover 21. The rocket leaves the launcher with a forward velocity of approximately 2.20 feet per second and an acceleration of approximately 36 0 gs. Over the range for which this rocket is designed to operate, the effect of a crosswind is self-compensating. The cross drift effecting the rocket at its initial low velocity about equals the action of the rocket in heading into the wind during its latter high velocity flight to its target. The configuration of the nozzle cover 21 is such that the rocket will have this required stability, that its yaw will be minimal, and that the rocket axis and the axis of the punch 16 will be tangential to the line of flight so that upon impact the punch will be properly presented to the target.

The curves of FIG. 6 depict the speed of the rocket, curve 72; its time of flight, curve 73; and its altitude during flight, curve 74. These curves are illustrative of the short time between firing and impact on the target. The altitude curve shows that small errors in aiming will have little effect on the accuracy and probability of a hit within the intended range of use. The short time to maximum casing pressure and velocity means that the device is effective at short ranges.

The curves of FIG. 7 are plotted against an ordinate representing estimated, or design, gross weights of rockets. The velocity constant of 7,000 f.p.s., for curves 76 and 77, is close to the minimal velocity needed to obtain the desired flat trajectory and the desired time between firing and impact to reduce errors caused by movement of the target so that direct sighting of the target may be used. This minimal velocity is about 5,000 feet per second. It is, also, the minimal velocity which will give, over the indicated range, the desired minimal impact pressures. While penetration is a function of slenderness and density, it is, also, a direct function of velocity but increases in velocity result in a geometric increase in the total weight of the rocket as is evident from curve 75 of FIG. 7, a plot of ordinal estimated gross weights against velocity. Thus, the velocity must be kept to that minimum which will give the needed penetration if the desired low rocket weight is to be achieved. According to curve 75, this minimum is about 5,000 f.p.s. Curve 76 depicts rocket weight versus density of the punch load as the abscissa where the punch has a slenderness ratio of eight, an expected penetration of 12. inches, and a velocity of 7,000 feet per second. This curve indicates that, with these constants, to be able to achieve a light weight rocket that can be fired by a single man, punches of a density greater than 0.5 pound per cubic inch must be used. Curve 77 shows the requirement for slenderness ratios of about ten and much greater. The abscissa of this curve are values of the ratio of the length of the punch to its diameter with the same values of velocity and expected penetration as for curve 76. Thus, the curve 77 is a curve of constant momentum per unit cross-sectional area of the punches having such slenderness.

Projectiles having a slenderness between 7 and 10 have been fired by the Ballistic Research Laboratory at the Aberdeen Proving Grounds as reported in Ordnance Corp Pamphlet, ORDP -245, May 1957, pp. 2-119, and the results of these shots at 60 obliquity to the target normal are shown in FIG. 8 by points A and B which are plots of ballistic limit against penetration (normal to face) over core diameter, or punch diameter. These punches were formed from steel. The other points, C to I, inclusive, on this graph represent the test data from shots made by the above BRL at applicants request to check the concept of the herein disclosed invention. The target consisted of two parallel plates of two-inch and four-inch thick homogenous armor, separated by a six-inch air gap. These shots were, also, made at sixty degrees to the armor normal but were made at higher velocities and with punches of greater density and slenderness. From this figure, it will be seen that there is poor or no correlation between the prior art and the present disclosure. Penetrations have been achieved which were not predictable for increased penetration should lie on the lines through A and B and thereabove. On these lines, penetration is a function of velocity. The test data for points C to I, inclusive, is given in the following table:

punch, and from the case to the disc and then to the punch, upon impact of the punch on the target.

In the explosive field, penetration of armor by means of cavity-lined shaped charges is considered to be probably the most efiicacious method, yet, penetration achieved as indicated hereinbefore approximate those for an idealized jet where the velocity is above 20,000 feet per second. The penetrations depicted in FIG. 8 by the points C to I, inclusive, are within seven to twenty percent of those which would be obtained by an idealized jet where the penetration is a function of the punch length times the square root of the ratio of punch and target densities. See: The Science of High Explosives by Melvin A. Cook, page 252, published 1958 by Reinhold Publishing Corporation, New York, N.Y. The above idealized formula for jet penetration, when once appreciated as being applicable to a high velocity punch, indicates the need for a punch that has a greater density than the target and, also, it indicates the probable length of punch that must be used to penetrate a given piece of armor. However, the above reference to shaped charge phenomena is not a statement of the theory of penetration found in the present invention. It is only a comparison of results. Further, shaped charge phenomena does not teach the obtaining of such penetrations at the velocities nor by the means disclosed herein.

Given the structures set forth herein which it may be desired to penetrate, such as heavy armor, a device is hereby disclosed wherein a punch may be launched at distances up to two miles from such structures by a single person using a device weighing less than twenty pounds, and which can be directly sighted and fired, and which punch will penetrate such structures.

It is to be understood that the form of this invention, herewith shown and described, is to be taken as a preferred example of the same, and that various changes in the shape, size and arrangement of parts may be resorted to, without departing from the spirit of this invention, or the scope of the subjoined claims.

We claim:

1. A missile system including: an open ended launching tube having a uniformly smooth inner cylindrical surface; a rocket in said tube; rotating segmented sabot means having an outer cylindrical surface conforming to the inner cylindrical surface of said launching tube and an inner surface conforming to and extending along at least half of the outer peripheral surface of said rocket, said sabot means having segments of equal size and shape to support and rotatably mount said rocket in said launching tube for rotation of said rocket as said rocket is being fired from said launching tube; and means for rotating said rocket relative to said launching tube as said rocket is fired from said launching tube.

2. A missile system as set forth in claim 1 wherein said sabot means has three segments, said segments having an inwardly tapered inside surface near one end that con- Diam- Pene P cos Length, eter, Density, tration, Velocity, Weight inches inches L/D lb./in. inches D ftJsec. lb.

forms to a nose cone portion of said rocket and an outwardly tapered inside surface near the other end that conforms to a tapered outer portion of a nozzle for said rocket, and said segments are adapted to be centrifugally removed from said rocket as said rocket leaves said launching tube.

3. A missile system as set forth in claim 1 wherein said the rocket case and the punch mounting disc 26 to the 7 means for rotating said rocket includes said rocket having a turbine secured thereto for rotating said rocket as said rocket is being fired from said launching tube.

4. A missile system as set forth in claim 1 wherein said rocket includes a body which is made up of a liner of substantially uniform thickness to render said liner capable of withstanding the longitudinal stresses and about half of the circumferential stresses exerted by pressure from Within said liner, and wrapping material means mounted on said liner so as to render said body capable of withstanding all the circumferential stresses.

5. missile system as set forth in claim 4 wherein said liner is made of aluminum and said wrapping material means includes high strength fiberglass cords.

6. A missile system as set forth in claim 1 wherein said means for rotating said rocket includes said rocket having a nozzle at one end with a spin turbine secured coaxially therewith, said spin turbine having a centrally located exhaust passage with a plurality of turbine blades projecting into the exhaust passage to be in the path of exhaust gases from said rocket nozzle exhaust, to cause rotation of said spin turbine and thereby said rocket and nozzle.

7. A missile system as set forth in claim 6 wherein said nozzle and spin turbine are secured together by coacting oppoled end edges of said nozzle and spin turbine with each of said nozzle and spin turbine having a circumferential groove in the outer periphery thereof and adjacent said opposed end edges, a segmented ring positioned around said opposed end edges and having a portion thereof fit'ted in each circumferential groove to secure said spin turbine to said nozzle, and means keying said nozzle and spin turbine for simultaneous rotation so long as said segmented ring is positioned around said opposed end edges.

8. A missile system as set forth in claim 7 wherein the outside diameter of said segmented ring, and the outside diameter of said spin turbine are in sliding engagement with said inner cylindrical surface of said launching tube.

9. A rocket having a drive nozzle, a spin turbine adapted to be secured to and coaxial of said nozzle along opposed end edges of said nozzle and turbine, each of said nozzle and turbine having a circumferential groove in the outer periphery thereof and adjacent said opposed end edges, a segmented ring positioned around said opposed end edges and having a portion thereof fitted in each circumferential groove to secure said turbine to said nozzle, means interconnecting said nozzle and turbine together for simultaneous rotation so long as said segmented ring is positioned around said opposed end edges, and said rocket being rotatably and slidably mounted relative to an inner cylindrical surface of an open ended launching tube in such a manner that the outer surface of said segmented ring and the outside surface of said turbine rotatably and slidably engage said inner cylindrical surface.

10. A device as set forth in claim 9- wherein said rocket is also rotatably and slidably mounted in said launching tube by rotating segmented sabot means that has an outer peripheral surface rotatably and slidably engaging said inner cylindrical surface and an inner surface that engages and extends along a portion of the outer peripheral surface of said rocket.

References Cited UNITED STATES PATENTS 2,775,163 12/1956 Vegren 89-1.808 2,939,275 6/ 1960 Loedding -255 2,945,421 7/ 1960 Pion 891.80'8 2,994,274 8/ 1961 Dunlap 10293 2,995,011 8/1961 Kimmel 60'255 3,015,991 1/1962 Forbes 891.816 3,089,388 5/1963 Webster et al. 891.816 3,033,116 5/ 1962 Critcher et al. 10293 SAMUEL W. ENGLE, Primary Examiner US. Cl. X.R. 

